(1) Field of the Invention
The present invention relates to a reduced noise and performance improved rotor blade for a helicopter, particularly a reduced noise and performance improved main rotor blade of an economic helicopter with the features of the preamble of claim 1.
(2) Description of Related Art
It is well known that the noise of helicopters in descent flight is dominated by the blade vortex interaction noise. The blade vortex interaction noise occurs when the rotor blade interacts with vortices generated by previous blades.
The document U.S. Pat. No. 5,035,577 proposes a non-linear twist of the blade which consists in giving the outboard end of the blade, for example between 85 and 100% of the total radius of the blade R, an extra amount of twist. This has the effect of reducing the strength of the marginal vortex, or even of cancelling it, for a given amount of lift, so that the low-speed performance is improved and the noises of blade/vortex interactions during descent are attenuated. However, this arrangement does not make it possible to push back the limits of stalling, and the power savings decrease at high speed.
The document U.S. Pat. No. 5,332,362 proposes a radical increase in the amount of twist in a central region lying approximately between 45 and 80% of the span. A modification of this kind is aimed at improving the lift capability.
The document US 2005158175 A1 discloses a rotary wing blade of limited taper ratio, comprising in succession along a reference radius: an inner zone; a forward-swept zone; and a back-swept zone, where a zone is said to be swept “forwards” when its leading edge forms a positive angle relative to the axis of the blade, i.e. extends forwards in the direction of rotation of the rotor, and a zone is said to be swept “back” when its leading edge forms a negative angle. The beginning of said forward-swept zone is in the range 0.47 to 0.65 times the reference radius R of the rotary wing. The leading edge of the blade at the reference radius presents an angle gamma lying in the range −40° and −60° relative to a general axis of the blade.
The document U.S. Pat. No. 6,000,911 A discloses a rotor blade with an anhedral at its swept-back tip, said blade having a leading edge and a trailing edge and being formed of successive elemental cross sections identified by the distance r separating each of them from an axis of rotation, each having a defined chord profile and a center of pressure whose offset from the pitch-change axis, orthogonal to each of said sections, determines the sweep of said blade. Said blade is subdivided along its longitudinal length into four regions and its chord length L increases in a more or less linear manner in a first region, is at its maximum and constant in a second region, decreases linearly in a third region and decreases according to a parabolic function in a fourth region.
The document U.S. Pat. No. 6,116,857 A discloses a blade with a reduced sound signature for a helicopter main rotor and said document discusses the interaction of vortexes and rotor blades, generating an impulsive noise which can be very intense. The geometry of the rotor blade of said state of the art is optimised in order to reduce the noise emitted by a helicopter main rotor during certain flight phases, particularly in descent and in landing approach, phases of flight during which a significant noise source comes from the interaction of the rotor blades with the vortexes which they generate. Said disclosure is an integral part of the present application. Amongst other known means the document U.S. Pat. No. 6,116,857 A proposes for modifying the characteristics of the vortex at emission first known passive means consisting in slimming down the tip or the blade end, so as to displace the maximum local circulation toward the inside of the rotor, and different blade end shapes slimmed down on the chord have been proposed. A second known passive means according to the document U.S. Pat. No. 6,116,857 A consists in applying to the blade a twisting law leading to a weak blade tip circulation gradient, at the emission azimuths and another known passive means consist in adding a vertical aileron on the end profile of the blade tip in order to prevent or disrupt the rolling up of the vortex, or to add a spoiler on the leading edge at the blade end, in order to increase the viscous radius of the vortex. The document U.S. Pat. No. 6,116,857 A proposes a blade tip with a parabolic shape in plan.
The document EP 0482788 A1 discloses a flow separation behind the notch region of a swept tip of a helicopter rotor blade being reduced by features of the geometry and aerodynamic characteristics of the notch region itself. A forwardly swept leading edge portion extends at an angle between 30 degrees and 55 degrees from a reference line parallel to a blade pitch change axis and the leading edge of the aerofoil in the notch region incorporates blade droop.
The document EP 0867363 A2 discloses a root end attached to the rotor head for rotationally driving. The central portion has aerodynamic characteristics depending on the leading and trailing edges and which extend linearly from the root end in parallel to each other, and the chord dimension (C) therebetween. A planform shape of the blade tip portion is defined by the first leading edge which extends forwardly as the distance from the outboard end (P1) of the leading edge of the central portion outwardly increases, the second leading edge and the side edge which are rearwardly swept as the distance from the outboard end (P) of the first leading edge toward outboard side outwardly increases, the first trailing edge which is curved forwardly as the distance from an outboard end (P5) of the trailing edge of the central portion outwardly increases, and the second trailing edge which is swept rearwardly as the distance from the outboard end point (P6) of the first trailing edge outwardly increases. This configuration makes it possible to eliminate the delocalization in the supersonic region and greatly reduce high-speed impulsive noises.
The document GB 1247966 A discloses a rotatable airfoil with an inboard portion and an outboard swept portion which has a concave forward swept portion and a convex aft swept portion, the cosine of the acute sweep angle at each point along the outboard swept portion being inversely proportional to the radial distance r of that point from the axis of rotation. In a second embodiment, there is provided an additional outboard aft swept portion with a sweep angle of substantially 70 degrees. The leading edge of the airfoil is shaped to provide a corresponding angle of sweep in the line of minimum air pressure which is substantially parallel to the leading edge. The forward and aft sweeps delay the onset of “compressibility” problems and the additional sweep reduces tip vortex effect.
The document JP 2002308192 A discloses a rotor blade with an inner blade part having a front edge and a rear edge linearly extending in parallel from a base end part along a feathering axis F, an advance part having a parallel linear front edge and rear edge displacing frontward from the outer end of the front edge and the rear edge of the inner blade part respectively, and a retreat part having a parallel linear front edge and rear edge displacing rearward from the outer end of the front edge and the rear edge of the advance part respectively. A blade area Sf surrounded with the extension line of a ¼ chord line of the inner blade part, a ¼ chord line of the advance part, and a ¼ chord line of the retreat part and a blade area Sr surrounded with the extension line of the ¼ chord line of the inner blade part and the ¼ chord line of the retreat part are set to approximately equal to each other.
The document WO 2008091299 A2 discloses blades for rotorcraft designed and/or implemented with a swept portion that occupies at least 20-40% of a length of the blade. Forward and aft sweeps are contemplated, with up to 20 DEG or more of sweep. The swept portion preferably has a thickness ration of at least 10-20% at R80, and can have a tapered planform with a relatively outboard section having a smaller chord than a relatively inboard section. Contemplated design methods include optimizing or otherwise designing the rotor blade planform and lift distribution along the blade for efficiency in various flight conditions without taking into account the detrimental effects of high Mach numbers, and then using sweep angle, airfoil thickness and transonic airfoil shaping to maintain the lift distribution, low drag and low noise level at real Mach numbers at the various blade stations at the various flight conditions.
The document WO 2008147376 A1 discloses rotor blades pre-bent in at least one of a flap direction and a lag direction, wherein the pre-bent portion comprises at least 20-60% of the length of the blade. Preferred methods include analyzing the rotor dynamic behavior using computational methods, deciding on the operational case (rotor lift load, forward speed, etc.) in which the loads and vibration reductions are desired, and using the computed results to decide on an amount of pre-bending of the unloaded blade so that it comes closer to the feather axis under load. Another class of preferred methods models the bending of a first blade in flight loading conditions, and then designs a second blade having a pre-bend in approximately an equal in magnitude and opposite in direction to the bending. It is contemplated that such “pre-bent” blades can significantly reduce rotor loads and vibration levels of rotorcraft equipped with semi-rigid or rigid rotors.
The document US 2006269418 A1 discloses a main rotor blade exhibiting a unique planform shape in which the blade chord increases from the root end of the blade inboard region to the outer main region of the blade, where the blade chord achieves a maximum chord at a spanwise location within the main region, then decreases toward a distal tip end. The leading edge preferably is generally straight while the trailing edge is contoured to define the chord. Another characteristic feature of the rotor blade design is the location of the blade-feathering axis in which the feathering axis is located at a mid chord position over some inboard length of the rotor blade then transitions to a quarter chord location. Another characteristic feature is an airfoil distribution along the blade span that transitions from a blunt trailing edge to a sharp trailing edge airfoil suited for mid-range Mach number operation. The tip region preferably utilizes a transonic flow airfoil. Another characteristic feature of the rotor blade design is an unconventional combination of positive and negative twist gradients.
The document U.S. Pat. No. 3,822,105 A discloses a helicopter blade shaped to have a tip of selected twist, camber, planform, thickness distribution, sweep and airfoil so as to increase rotor operating efficiency, reduce rotor noise, and to reduce or eliminate rotor instability.